Aircraft proportional navigation



Dec. 14, 1965 K. BRUECKER-STEINKUHL 3,223,357

AIRCRAFT PROPORTIONAL NAVIGATION Filed June 12, 1962 2 Sheets-Sheet 1ObJecf/I/e l l l I J A/rcraff fimemafics objecfive g I F I I l A/rcraffF" l 3 Noumea/1 speeif p n be/mv/ar I 2 k L I l fFinemaf/as INVENTOR KUR T BRUECKER-TEINKUHL BYW ATTORNEYS De 1965 K. BRUECKER-STEINKUHL3,223,357

AIRCRAFT PRQPORTIONAL NAVIGATION Filed June 12, 1962 2 Sheets-Sheet 2Objecfiv I I I A/rcraff v); Noumea/7. Rudder Spec/fie Q N confro/behavior Y I C D- mvc/Ian/lr F6) I klhemaf/bs Fig.3

Fig.4 4

INVENTOR KURT BRUECKER-S TE/NKUHL ATTORNEYS United States Patent 03,223,357 AIRCRAFT PROPORTIONAL NAVIGATION Kurt Bruecker-Steinkuhl,Liesegangstr. 10, Dusseldorf, Germany Filed June 12, 1962, Ser. No.202,943 Claims priority, application Germany, June 19, 1961, B 62,953;June 25, 1961, R 63,393 3 Claims. (Cl. 24414) The purpose of thisinvention is to facilitate the guiding of aircraft toward movingobjectives, and especially to introduce certain improvements into themethod usually known as proportional navigation so as to adequatelycompensate for the aerodynainical and technical disturbances which arecommonly encountered under actual flying conditions.

In the drawings:

FIGURE 1 is a simplified diagrammatic view showing only the kinematicsof aircraft steering;

FIGURE 2 shows diagrammatically a navigation system in which theaerodynamic influences are taken into consideration;

FIGURE 3 shows the improved system and its mode of operation inaccordance with this invention; and

FIGURE 4 shows diagrammatically the essential parts of the controlapparatus in their operative relationship to each other.

Proportional navigation is defined as a method of steer ing in which theangular velocity of the flight direction vector of the aircraft isproportional to the angular veloc ity of the line of sight from theaircraft to its objective. According to the usual and known method ofproportional navigation, the angular velocity 6;; of the line of sightis measured, and is then multiplied in a navigation mechanism with thefactor k, the so-called navigation constant. According to A. S. Locke,Guidance, Princeton, NJ., 1955, p. 475, the expressions in thespecification angular velocity of the line of sigh and angular velocityof the flight-direction vector are also physically equivalent to theexpressions rate of rotation of the line-of-sight from the missile tothe target and rate of change of missile heading. In the ideal case setforth in FIGURE 1, it is assumed that the angular velocity of the flightdirection vector can really be kept equal to the expression kgb.

The simplified version of FIGURE 1 is valid only in a case whereconsideration is given solely to the kinematics of the steering process.If other aerodynamical and technical influences are taken intoconsideration, then it will .be found that the angular velocity 'y ofthe flight direction vector is not at all equal to la but insteadfollows with a certain retardation the prescribed and variable commandvalue. Such retardation is caused first by the aerodynamic or specificbehavior of the aircraft, and second by the retarding members of thecontrolling and setting mechanism.

FIGURE 2 sets forth the method of proportional navigation withconsideration given to the aerodynamic in; flucnces. The angularvelocity 3 of the line of sight, which is picked up by theobject-seeking head (a parabolic antenna), is multiplied in thenavigation mechanism by the constant k and is then fed into the steeringcontrol mechanism with a steering adjustment '7]. In doing so, it isassumed that the steering adjustment 1 is proportional to the fed-invalue k qb. The angular velocity 6 of the flight direction vector doesnot instantaneously follow the steering adjustment 1;, but because ofthe aerodynamic influences, will follow only with a certain delay,designated as specific behavior in the block in FIGURE 2. The expressionspecific behavior relates to the aerodynamic behavior of the airship.When the rudder is repositioned the airship will not immediately enterupon the course corresponding to the repositioned rudder but only "iceafter some delay. This delay is determined by aerodynamic influences.The angular velocity j of the flight direction vector changes in turnthe angular velocity {0 of the line of sight, as has been indicated byfeedback line designated kinematic on FIGURE 2. The new angular velocitygb of the line of sight, which is dependent not only on "y but also on"5/ (the angular velocity of the line of flight vector of theobjective), will, as in the previous case, be taken over by theobjective-seeking head, for processing in the airship. In the closedcontrol circuit of FIGURE 2, the angular velocity (b of the line ofsight corresponds to the magnitude of regulation, and the steeringdisplacement 1 to that of the setting. The navigation constant k in theideal case of FIGURE 1 corresponds here to a navigation factor whichcomprises a constant k which is independent of the frequency, and afactor 1 C2.S +C .S+1 which is, on the other hand, dependent on the saidfrequency.

By S. S. Chin, Missile Configuration Design, New York, 1961, p. 147, theaerodynamic transfer function is given as The formula given in theapplication agrees with the formula of Chin if the aerodynamic staticgain K/w =1 and if 6:1 Also C =1/w and C =2w where w is the undampednatural frequency of the missile and g is the damping constant. Theexpression F(S) depends upon (S) =5+jw*. It is thus dependent upon thefrequency. On the other hand the transfer function 1 =k (see FIG. 2) isindependent of the frequency. That means that the rudder displacement 17can be kept proportional to the measured angular velocity A diflicultywith the usual method of proportional navigation described above is thatthe stability of the method is relatively small, and especially whenthere is only a short distance between the aircraft and the objective.Investigations have shown that there is an upper limit of stability forthe navigation factor constant, which may not be exceeded, for otherwisedisturbing vibrations would be set up, and the aircraft would lose itsobjective. This upper limit of stability becomes less as the distance ofthe aircraft from the objective decreases. Hence because of stabilityconsiderations the navigation factor constant may not be given a veryhigh value. For other technical reasons, to keep the line of flight asstraight as possible, and to approach the constant bearing course, andto keep the transverse acceleration of the aircraft low, it wouldnevertheless be desirable to keep the navigation factor constant as highas possible.

The method according to this invention is therefore characterized inthat the stability of the steering method, as compared with the usualsteering methods, for the same value of the navigation factor constant,is greatly improved, so that the navigation factor constant may then beincreased. It will therefore be possible to navigate the aircraft withadequate stability and thus increasethe closeness of safe approach ofthe aircraft to its objective.

According to this invention, for steering the aircraft, differentangular velocities are measured, namely that of the line of sight fromthe aircraft to its objective, and that of flight-direction vector ofthe aircraft.

The individual angular velocities are measured in the following manneras disclosed by A. S. Locke, Guidance, Princeton, N.J., 1955, pages350-353:

(a) In the aircraft the antenna of the search head is rotatablysuspended from a gyroscope and is therefore firmly held in the directionof the target. Every aberration of the line of sight introduces an error.9, which is converted into a command for adjustment of the antenna. Therequired antenna movement 2A to compensate for this error is thenmeasured by means of a rate gyro. 5A is equal to the angular velocity(,2: of the line of sight.

(b) The angle 7 between the direction of flight and a stationary axis ofreference is measured continually by a wind vane. By changing thedirection of flight the velocity ''y is determined directly in theairship according to one of the known constructions by means of a rategyro. The relationships of the separate steps of this process areexemplified in FIGURE 3. The angular velocity of the line of sightmultiplied by the constant k represents a command value, with which theactual angular velocity y of the line of flight vector is compared atthe place N. The difference between these two values, namely between thecommand value and the actual value of the angular velocity of the lineof flight vector, is fed 'into the steering control mechanism as asupplementary steering adjustment. Here it is again assumed that thesteering adjustment 1; is proportional to the fed-in value, which inthis case is equal to a dilference between angular velocities.

The fed-in value is the value (k- ;o"y) at the input of the ruddercommand mechanism which controls the rudder displacement 1;, the latterbeing proportional to this input value, the proportionality factor beingk According to FIGURE 3, the measurement of the angular velocity of thedirection of flight vector involves carrying back this angular velocityto the subtraction place N, as for instance with a feedback coupling.The feedback circuit that is inside and to the right of the aircraft inFIGURE 3 represents a supplementary control circuit, and the steeringmethod according to this invention consists in this: that the angularvelocity in the supplementary regulating circuit is fitted as closely aspossible to the variable command values k'rp (angular velocities of theline of sight 2, multiplied by k) before the angular velocity 'v in themain regulating circuit influences the angular velocity of the line ofsight (2: over the kinematic feedback. The command values of the maincontrol circuit are influenced by angular changes in the line of sight,and therefore in the position of the objective, as well as by thereaction from the supplementary as well as a factor that is dependent ofthe frequency in which the values C and C corresponding to the feedback,are strongly distinguished from the values C and C By navigation factoris understood the static amplification of the open regulating circuit.The static amplification of the regulating circuit shown in FIGS. 1-3 istherefore given by (Fig. 2) 1 =IC -m kg f0! (Fig. 3)

The following relations will'therefore hold true:

This difference influences stability in such a manner that the upperlimit of stability for the navigation factor constant, with thisinvention is :much higher than with the usual steering methods ofproportional navigation. The navigation factor constant or thestrengthening of the navigating mechanism according to this invention isso chosen that the upper limit of stability of the constant will not beexceeded until the aircraft is only a short distance from its objective,and that up to that point the airship will be stable during steering.Further explanation of the above formulas is as follows: At the mixingpoint N in FIG. 3 the ingoing and outgoing values are given by theformula:

from which the transfer function is obtained as follows:

The equations for the regulating circuit can be set up from FIGS. 2 and3 either with or without the supplementary regulating circuit, wherebyapproximate expressions will be used for the kinematics. As is customarywith such regulating circuit calculations, the necessary conditions forstability for the strengthening or for existing navigation factor can beset up with the help of the stability criterion of Hurwitz. Afterextensive calculation, the following condition is arrived at for theregulating circuitof FIG. 2:

and for the navigation factor of the regulating circuit with thesupplementary regulating circuit of FIG. 3, the following conditionwith 1) cos 6 and where r=distance of airship from objective v=ve1ocityof airship v =velocity of objective the p v cos 6 the upper limit ofstability is approximately proportional to the distance r. Hence for acertain value of between about 3 to 5, the upper limit is exceeded onlyfor low values of r. Instabilities and disturbing oscillations in theguiding circuit can then have only little effect, and the distance ofthe airship from the objective while passing the latter, namely, themiss-distance, is much reduced. All of these results have been confirmedby electronic computers with actual numerical trajectory calculations.

Since differences between the command and actual values of the angularvelocity of the flight direction vector are promptly received by thesupplementary control circuit and used for influencing the aircraft, theelimination of a disturbance or of a deviation from the command valuecan follow more quickly and more directly. Investigations have shownthat in consequence thereof the line of flight under adverse conditionswill suffer less disturbance, and that the miss distance will be muchreduced. The advantage of the invention therefore consists briefly inthis, that there is improved stability during steering, that there isless disturbance of the line of flight, that the missdistance isreduced, and that the chances of reaching the objective are muchimproved.

The measurement of angular velocities is sometimes more difficult thanthe measurement of angles. It may therefore be preferable to use angularmeasurements rather than angular velocity measurements. In the knownprocess of proportional navigation, the angular velocity of the line ofsight is measured with gyroscopic devices as shown by A. S. Locke,Guidance, Princeton, N.J., 1955, pages 350-353. The measurement of theangular velocity of the flight path vector of the aircraft is made in asimilar manner. Another feature of this invention is that measurementsare first made of the angles which the line of sight and the line offlight make with a stationary reference line, and the difference betweenthese two angles is then delivered to the steering command mechanism tohelp regulate the same. At the same time this angular difference isdifferentiated in a differentiator which is associated with the steeringcommand mechanism, and is therein converted into the desired angularvelocity difference. As stated above, continuous measurements can bemade of angle (,0 of the line of sight by the trailing of the antenna,and also of the angle 7 by the flight direction indicator. Aftertranslation of the angular measurements inotelectric voltages, thedifference (k- -'y) and in an electric differentiator of knownconstruction the value (k- &) namely, the angular velocity difference,is formed. This subsequent differentiation of the angular velocitydifference has the advantage in that initially it is only the angles andnot the angular velocities that need to be measured, and that thedifferentiation can thereafter he carried out in a fixed network whichin some cases may be formed entirely of resistors and capacitors. As astationary line of reference, against which the angular positions of theline of sight and the line of flight are measured, it is preferable touse a line that is specially provided for that purpose. The stationaryline of reference is a line in a fixed position in space. It can passfor example through a gyroscopic axis.

The angular velocity (0 which is measured in the manner above describedis converted into an electric voltage and is delivered to an electricamplifier. By means of the amplification regulator of this navigationamplifier the value k is suitably fixed and introduced.

In accordance with the lowering of the limit of stability for thenavigation factor, such' factor can be higher at greater distances ofthe aircraft from its objective than at smaller distances. According toa further feature of this invention the amplification of the navigationmechanism (factor k) is not maintained constant but changes duringflight, especially during the lowering of the upper limit of stabilitythe amplification is reduced. The navigation factor of this invention isgiven by This expression will change in value when, with k remainingconstant, the factor k changes; k is equivalent to the electricamplification of the navigation mechanism. This amplification iscontrolled as set forth below.

The changing or controlling of the amplification can conform to aprescribed program, according to which the upper stability limit duringthe normal flight procedure will be surpassed only after close approachof the aircraft to its objective. The changing or controlling of theamplification does not always need to follow a fixed program, but can bevaried according to a measurable factor. It is especially advantageousaccording to this invention, to measure, during flight, the distance ofthe aircraft from the objective, and to regulate the amplification inaccordance with this measurement. The measurement of the distance of theairship from the objective is preferably made by determination of thetime of travel of an electric impulse which is radiated from the airshipand after reflection from the objective is received by the airship byknown means. The advantage of such a measurement is that theamplification can then be kept as high as possible, but always withinthe limits of stability. In this manner the aircraft can be steered withstability as close as possible to the objective.

The upper limit of stability for the amplification or the navigationfactor constants, is inversely proportional to the speed of theaircraft. Hence to additionally increase the stability of the process atsmaller distances, provision is made for reducing the aircraft speedupon close ap proach to its objective. The velocity of the airship canbe reduced by the release of braking surfaces positioned in the path ofthe air current by known means. This reduction of speed, as well as acorresponding change of amplification, can be effected either inaccordance with a prescribed time-controlled program or according to avariable factor, especially to the measured distance of the aircraftfrom its objective.

In FIGURE 2 the adjustment 1; and also the angular velocity "y isproportional, up to a certain limit, to the adjustment 4;, whereas inproportional navigation there is a proportional regulation whichinherently introduces a continual error. The result of this is that thevelocity of the line of sight cannot be returned to its command valuezero during continual maneuvering of the objective, but will receive aresidual deviation. Ely continually changing the angular direction ofthe objective, by maneuvering the line of sight will also change itsposition in space. The transverse acceleration of the aircraft will besmaller as the deviation of the velocity of the line of sight from thecommand value zero is smaller. Since the transcommand mechanism forsteering adjustment.

' factor k verse acceleration to which the aircraft subjects itself mustnot exceed a certain value, it is therefore essential that any residualdeviation should be kept as small as possible.

In proportional regulation the residual deviation becomes smaller whenthe static amplification; or in the case of proportional navigation, thenavigation factor; is increased. Nevertheless there are limits to theincrease of static amplification or of the navigation factor, since withincreasing amplification the yawing tendency and the danger of exceedingthe upper limit of stability are increased.

According to another feature of this invention it is advantageouslypossible, without any increase of the navigation factor, to completelyeliminate any residual deviation of the steering process during uniformmotion. For this purpose, after measurement of the angular velocity ofthe line of sight, there is produced in the navigation mechanism alsothe integral value of this angular velocity.

Both values, namely the measured angular velocity of the line of sightand the corresponding integral value, are multiplied with differentsuitably selected constants, and thus adequately amplified. In the usualproportional navigation the measured angular velocity j is multiplied bythe factor k, which means that an electric signal of the value kisproduced and delivered to the rudder command mechanism. In accordancewith this invention, after measurement of 2, in the navigation mechanismthe integral f dt is also formed and multiplied by a second factor k sothat there will then be produced a second signal of the magnitude k fbdt. The sum of the two signals (k- +k -fdt) is delivered to the ruddercommand mechanism. The factor k is chosen about equal to k. Thesummation value of both of these adequately amplified values is thenproduced, and is delivered to the The essential feature here is theintegration of the angular velocity of the line of sight, and theimproved method can therefore be designated as Proportional IntegralNavigation.

It may also be advantageous, for the purpose of stability, aftermeasurement of the angular velocity of the line of sight, to produce inthe navigation mechanism also the differential quotient of this angularvelocity, to be used for steering the aircraft. In this case thesummation value is produced in the navigation mechanism from (1) theangular velocity of the line of sight multiplied by a first constant,and (2) the differential quotient multiplied by a second constant. Thesummation value is fed into the mixer station of the steering commandmechanism. After measurement of (10, the differential quotient of 1,namely d i/dt=q' is formed and is multiplied by a second The sum of thetwo signals (kb-j-k is delivered to the rudder command mechanism.

Alternatively the summation value which is to be used for influencingthe steering command mechanism may also be produced in the navigationmechanism from three values, namely (1) the angular velocity of the lineof sight multiplied by a first constant, (2) the integral valuemultiplied by a second constant, and (3) the differential quotientmultiplied by a third constant. As already mentioned, the measuredangular velocity {u is converted by the navigation mechanism into anelectric voltage. From the continuously adjusted voltage the integratedvalue of 4: or the differential quotient of can be formed by an electricintegrator or differentiator of known construction. The integral valuesor differential quotients thus obtained are again converted intoelectric voltages whichare strengthened in subsequent amplifiers ormultiplied by a constant. The various voltages can be added in any knownprescribed manner.

The constants or the amplifications, of integral value,

and of the differential quotient, are chosen in such a manner that theupper limits of stability of these constants will not be exceeded up toclose approach to the objective. Since the stability limits of theseconstants drop from high values to low values as the distance of theaircraft from its objective diminishes, it is furthermore contemplatedin this process to change these constants during flight, especially fromhigher values to lower values, which may be done either in accordancewith a time schedule or corresponding to the flight, and preferablyafter measurement of the distance of the aircraft from the objective.The factors k, k and k correspond to the amplifications by amplifiers towhich the values ,fdt, d S/dt are delivered. To change the values of thefactors k, k k it is only necessary to change the amplification of theseamplifiers by means of variable potentiometers.

The process of this invention is illustrated by the example in FIGURE 4which shows the essential elements of an aircraft which are necessaryfor performing the steering process of this invention. For betterclarification, all details which are not necessary for an understandingof this invention have been omitted.

In FIGURE 4 the objective seeking head or parabolic antenna 1 serves toreceive and measure the angular velocity 13 of the line of sight. Aspreviously explained, that angular velocity 4 is measured by thetrailing of the antenna. This measured angular velocity, after beingconverted into a corresponding electric voltage, is then fed into theamplifier 2 to be multiplied by an amplification constant k. The value kis then passed from 2 to the mixer station 3. The direction of flight ofthe aircraft 7 is then measured by a wind vane 4. From the currentlymeasured value a, the angular velocity is obtained by differentiation,and after being converted into an electric potential, is led into themixer station 3. In this mixer station 3 there is obtained thedifference between the command value kqi and the actual value whichdifference is then delivered to the amplifier 6, in which thisdifference is further amplified by the factor k The amplification k isproduced by adjustment of the variable amplification control 7. From theamplifier 6, the adequately amplified voltage difference is fed to theservomotor 8 for adjustment of the rudder 9, and the direction of flightof the aircraft thus controlled. Blocks 1 and 2 contain the navigationmechanism, and block 6 the steering command mechanism.

The method of this invention is however not to be limited by theillustrated examples.

What is claimed is:

1. In an aircraft having steering mechanism and servomotor means foradjusting the steering mechanism to direct the aircraft toward anobjective, a regulating apparatus on the aircraft for improving theproportional navigation of the aircraft comprising: means for measuringthe angular velocity of the line of sight from the aircraft to anobjective and for converting the measurement into a first electricpotential; means for measuring the angular velocity of the direction offlight of the aircraft and for converting the measurement into a secondelectrical potential; means for multiplying said first electricalpotential by an arbitrary constant to provide a command value; a mixerstation having means for measuring the difference between said secondelectrical potential and said command value; and means for amplifyingsaid difference and feeding it to said servomotor means for adjustingsaid steering mechanism.

2. In an aircraft having steering control mechanism and servomotor meansfor adjusting the steering mechanism to direct an aircraft toward anobjective, a regulating apparatus on the aircraft for improvingproportional navigation of the aircraft comprising: means for measuringthe angular velocity of the line of sight from the aircraft to anobjective and for converting the measurement into a first electricpotential; means for measuring the angular velocity of the direction offlight of the aircraft and for converting the measurement into a secondelectrical potential; means for determining the distance of the aircraftfrom its objective and for converting the measurement into a variablevalue; means for reducing said variable value as the aircraft approachesits objective, variable amplification means for multiplying said firstelectrical potential by said variable value to provide a command value;a mixer station having means for measuring the difference between saidsecond electrical potential and said command value; means for amplifyingsaid difference and feeding it to said servomotor means for adjustingsaid steering mechanism with the adjustment thereof being proportionalto said difference.

3. In an aircraft having steering mechanism and servomotor means foradjusting the steering mechanism to direct the aircraft toward anobjective, a regulating apparatus on the aircraft for improvingproportional navigation of the aircraft comprising: means forestablishing a reference line in a fixed position in space; means formeasuring the angle of the line of sight from the aircraft to an objectwith respect to said reference line and for converting the measurementinto a first electric potential; means for measuring the angle of thedirection of flight with respect to said reference line and forconverting the measurement into a second electric potential; means formultiplying said first electric potential by a constant factor toprovide a command value; means for differentiating the differencebetween said command value and said second electric potential to obtainan angular velocity difference; and variable amplification means formultiplying said angular velocity difference by a constant and forfeeding it to said servomotor.

References Cited by the Examiner UNITED STATES PATENTS 2,463,362 3/1949Doll 250-230 X 2,992,423 7/1961 Floyd et al. 24414 3,083,666 4/1963Agins 24414 BENJAMIN A. BORCHELT, Primary Examiner. SAMUEL FEINBERG,Examiner.

UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTION Patent No.3,223,357 December 14, 1965 Kurt Bruecker-Steinkuhl It is herebycertified that error appears in the above numbered patent requiringcorrection and that the said Letters Patent should read as correctedbelow.

Column 1, line 49, for "y" read line 50, for "kqv" read k9 column 25lines 17 and l, for that portion of the formula reading "C2.S read C 5lines 25 and 26, for that portion of the formula reading K K S +Zw readi column 3 lines 37 and 38 for "4" each occurrence, read d column 4lines 1 to 6 for that portion of the formula reading "k read k column 5,line 7, for

"v" read v line 8, for "6 read v column 7, line 28 for "Jk/dt" read fdtline 51, for second occurrence, read P line 52 for [k';'!+k2 'y)" read(k-q' k -a column 8 line 29 for "value read value Y Signed and sealedthis 20th day of December 1966.

(SEAL) Attest:

ERNEST W. SWIDER EDWARD J. BRENNER Attesting Officer Commissioner ofPatents

1. IN A AIRCRAFT HAVING STEERING MECHANISM AND SERVOMOTOR MEANS FORADJUSTING THE STEERING MECHANISM TO DIRECT THE AIRCRAFT TOWARD ANOBJECTIVE, A REGULATING APPARATUS ON THE AIRCRAFT FOR IMPROVING THEPROPORTIONAL NAVIGATION OF THE AIRCRAFT COMPRISING: MEANS FOR MEASURINGTHE ANGULAR VELOCITY OF THE LINE OF SIGHT FROM THE AIRCRAFT TO ANOBJECTIVE AND FOR COVERTING THE MEASUREMENT INTO A FIRST ELECTRICPOTENTIAL; MEANS FOR MEASURING THE ANGULAR VELOCITY O THE DIRECTION OFFLIGHT OF THE AIRCRAFT AND FOR CONVERTING THE DIRECTION OFFLIGHT OF THEAIRCRAFT TRICAL POTENTIAL; MEANS FOR MULTIPLYING SAID FIRST ELECTRICALPOTENTIAL BY AN ARBITRARY CONSTANT TO PROVIDE A COMMAND VALUE; A MIXERSTATION HAVING MEANS FOR MEASURING THE DIFFERENCE BETWEEN SAID SECONDELECTRICAL POTENTIAL AN SAID COMMAND VALUE; AND MEANS FOR AMPLIFYINGSAID DIFFERENCE AND FEEDING IT TO SAID SERVOMOTOR MEANS FOR ADJUSTINGSAID STEERING MECHANISM.